Heretofore, various types of cooling structures for gas turbine drive blades have been made public. In FIG. 4 is shown a typical prior art design for a cooling structure for the air-driven blades in a gas turbine. With such a cooling structure, the air which enters via channels 4a and 4b on blade base 1 flows into blade cooling channels 5a and 5b within blade 3 in the direction indicated by the arrows; in this way it cools blade 3.
The air which flows from channel 4a on blade base 1 into blade cooling channel 5a on the leading edge 3a of blade 3 traverses a number of fins 13 (turbulators). As it flows through blade cooling channel 5a, which winds back and forth to follow the shape of drive blade 3, the air cools drive blade 3. It then flows out via hole A on the thin tip 14 of the blade and is mixed in with the main gas flow.
The air which flows from channel 4b on blade base 1 into channel 5b on the rear half of the edge of blade 3 must pass back and forth around a number of fins 13 which are provided in channel 5b. The air cools the trailing edge 3b of the blade via pin fins 15, then flows out through holes or slits B to mix with the main gas flow. A number of drive blades with this sort of high-speed cooling configuration are placed adjacent to each other along the circumference of platform 16 and set into disk 17.
Devices of the prior art such as those described above have hollow drive blades with a configuration in the base of the blade or its interior to provide high-speed cooling. However, since the platform from which the cooling components protrude is not itself cooled, the cooling capacity is insufficient.
Although the drive blade platform of a high-temperature gas turbine must be cooled, cooling it effectively induces thermal stress which must then be mitigated. Temperature differentials in excess of 1,000.degree. C. may occur between the air in the gas seal on the side of the platform with the gas channels and the air in the seal on the underside where the rotor is.
To address this problem, a number of configurations have been suggested which can effectively cool the platform surface and at the same time mitigate the temperature differential between the upper and lower surfaces of the platform.
One of these configurations, suggested by the present inventors, is published in Japanese Patent Publication 7-332004 of the Japanese Patent Office. In this configuration, holes are provided at the ends of the enclosed air channels which radiate from the center of the platform. Vents formed from shaped film are also provided on the upper surface of the same air channels. With this design the enclosed air which flows over the bottom of the platform passes through the holes at the ends of the radii, enters the shaped film vents and spreads out over the top of the platform to cool it effectively. If slits are provided which extend from the holes in the air channels to the edge of the platform, the expansion and contraction of these slits will mitigate the thermal stress occasioned by the temperature differential between the top and bottom of the platform. The slits will also prevent the platform from expanding.
Another such configuration was suggested by the present inventors in Japanese Patent Publication 8-246802. In this configuration, air channels are provided into which air is supplied from the base of the blade of a gas turbine on either its underside or its topside. This air passes through the interior of the platform in the vicinity of the bottom of the blade and then flows on either side of the blade. It is released at the end of either the top or bottom of the blade. In this way the platform is cooled.
Each of these configurations has its good and bad points. Currently there is a demand that a turbine operate at an even higher temperature in order to boost its efficiency. It would also be advantageous if the configuration used to cool the turbine could be formed using simpler techniques. Thus there is a demand for an efficient cooling configuration which requires fewer production processes.